A. Field of the Invention
The invention relates to aircraft instrumentation and control and, more particularly, comprises a method and apparatus for improved attitude determination and navigation, and an improved integrated flight information and control system in connection therewith.
B. Prior Art
Safe control and navigation of an aircraft requires a continuous stream of relatively accurate information concerning the dynamics of the aircraft as it moves along its flight path from origin to destination. Information such as the airplane""s altitude, velocity, heading, attitude (including at least pitch and roll information), among other data, are important for piloting the aircraft. Extensive, highly accurate, and correspondingly expensive, instrumentation has been developed for commercial and military aircraft to fill the required need. Such instrumentation is far too expensive for smaller, private aircraft, and the latter therefore must make do with simpler, cruder navigation systems providing more limited, and commonly less accurate, navigation data.
In larger commercial and military aircraft, inertial navigation systems have commonly been the instrumentation of choice. Such systems rely on the ability of gyroscopes to maintain their orientation in space once initialized, and to provide relatively sensitive indications of accelerations tending to disturb that orientation. These systems typically have significant mass and bulk, are expensive to acquire, and require continued, frequently costly, calibration, maintenance and repair to ensure continued acceptable performance. Their use has therefore been confined largely to larger commercial and military aircraft.
The advent of satellite positioning systems, such as the Global Positioning System (GPS) established and maintained by the United States, or the GLONASS system established and maintained by Russia, offers the possibility of significantly reducing the mass and bulk of many present navigation systems, and possibly their cost as well. Navigation systems using these facilities rely on the measurement of phase differences in received radio signals from a number of satellites in order to determine the position in space of the receiver, and therefore the platform on which the receiver is carried, with respect to the satellites. Because the position and velocity of the satellites relative to earth at any given time is known, the position and velocity of the receiver with respect to an arbitrary earth-based reference system can be determined from measurements with respect to the satellites.
In addition to navigation functions, such systems can also be used to determine the attitude of the vehicle in which the system is mounted. Numerous attitude determination systems based on GPS measurements have been proposed and such systems take many forms. For example, U.S. Pat. No. 5,548,293 issued Aug. 20, 1996 to Clark E. Cohen and entitled xe2x80x9cSystem and Method for Generating Attitude Determinations Using GPSxe2x80x9d proposes the use of a multiplicity of antennas on a vehicle whose orientation with respect to a reference frame is to be determined. The antennas provide a multiplicity of baselines from which the orientation may be found. Multiple baselines are used in order to resolve the position ambiguity inherent in measurements from antennas typically separated by many meters resulting from the short wavelengths used in GPS signaling (on the order of 0.2 meters). The use of a number of antennas, of course, increases the cost of the system, as well as the cost of installation, and inhibits the application of such a system to small aircraft in particular. Further, because of the very short wavelength, significant errors are introduced in the measurement whenever the distance between the antennas changes, as it is susceptible to do in response to stresses imposed on the aircraft during flight.
Some systems, such as that described in U.S. Pat. No. 5,534,875 issued Jul. 9, 1996 to Debra Diefes et al., entitled xe2x80x9cAttitude Determining System for Use with Global Positioning Systemxe2x80x9d, utilize a single GPS receiver and antenna on board the vehicle, but use standard inclinometers to determine the pitch and roll of the vehicle platform. Such hybrid systems fail to make use of the capabilities of GPS for attitude determination.
Still other systems, such as that described in U.S. Pat. No. 5,451,963, issued Sep. 19, 1995 to Thomas A. Lempicke, entitled xe2x80x9cMethod and Apparatus for Determining Aircraft Bank Angle Based on Satellite Navigational Signalsxe2x80x9d, utilize a single on-board GPS system that determines certain attitude information, such as bank angle, only under conditions of level flight, thus precluding effective use of the system in arbitrary maneuvers such as climbing or descending turns in which accurate attitude information is often most essential, particularly in connection with takeoff and landing. Further, the system posits a mode of operation (determining bank angle as inversely proportional to aircraft speed) which is not explained and not achievable by anything described in the patent.
Still another GPS-based system is described in U.S. Pat. No. 5,406,489, issued Apr. 11, 1995 to LaMar K. Timothy et al., entitled xe2x80x9cInstrument for Measuring an Aircraft""s Roll, Pitch and Heading by Matching Position Changes Along Two Sets of Axesxe2x80x9d. This patent uses both a GPS receiver and a multiplicity of accelerometers oriented along the three aircraft body axes, respectively, to determine attitude and other navigation information. Again, the hybrid nature of the system increases its cost, complexity, and maintenance requirements.
A. Objects of the Invention
Accordingly, it is an object of the present invention to provide an inexpensive but relatively accurate and reliable attitude determination method and apparatus for aircraft navigation.
Another object of the invention is to provide a simple, low-cost navigation system for determining attitude (roll, pitch) information for small aircraft.
Further, it is an object of the invention to provide an inexpensive attitude determination system that is useful as a backup for more elaborate navigation instrumentation systems.
Still a further object of the invention is to provide a simple, relatively inexpensive attitude indicator.
Yet another object of the invention is to provide an improved, economical integrated flight information and control system for control and navigation of aircraft.
B. Brief Summary of the Invention
In accordance with the present invention, we provide a method and apparatus for readily and inexpensively determining and displaying flight path angle and roll angle of an aircraft despite its engagement in arbitrary, but balanced, maneuvers such as ascent or descent accompanied by banked turns, conditions which present problems for many navigation systems. Because of its simplicity, and its reliance on a single source of measurement data for the requisite input information, the system is extremely simple and inexpensive to construct, install, and maintain. It is particularly suited for installation and use in small aircraft, where the cost of more elaborate and more expensive systems considered essential for navigation on larger aircraft effectively preclude their acquisition and use. Further, the system is sufficiently accurate and reliable to be used as a supplemental system on larger aircraft for use in integrity checking, as well as for backup in the event of failure of the primary system. The present invention obviates the use of a multiplicity of receivers or antennas or supplemental orientation indicators, and is useful throughout the entire range of flight dynamics commonly encountered in air navigation.
In particular, in accordance with the preferred embodiment of the invention, we determine the flight path angle xcex3 and roll angle xcfx86s of an aircraft and display these as parameters to the pilot as a principal measure of the aircraft attitude at a given moment. For convenience of reference, the stability axis roll angle xcfx86s and flight path angle xcex3 will on occasion be referred to hereinafter as the xe2x80x9cpseudo-attitudexe2x80x9d parameters or, more simply, as xe2x80x9cthe pseudo-attitudexe2x80x9d, in contrast to the conventional attitude parameters based on a body axis, namely, roll xcfx86 and pitch xcex8. The roll angle xcfx86s (xe2x80x9cpseudo rollxe2x80x9d) is advantageously determined with respect to the stability axis of the aircraft, as opposed to the body axis, since determination of the stability axis roll angle requires no knowledge of the angle of attack, a parameter which is frequently not known or readily determinable with accuracy in small aircraft. Similarly, the flight path angle xcex3 or xe2x80x9cpseudo pitchxe2x80x9d is determined about the stability axis of the aircraft in order to present a direct indication of the path of the vehicle through space. This is in contrast to conventional attitude navigation systems which determine and display roll angle and pitch about the aircraft""s body axis. Although the body-axis and stability-axis roll angle are typically nearly equal to each other, this is not the case with pitch angle, which differs from flight path angle by the angle of attack. Thus, in conventional systems, when a pilot wishes to maintain the aircraft on a particular flight path angle, he/she can not navigate by the pitch angle alone, but must correct it by the current angle of attack, a value which is frequently known only imprecisely at best in small aircraft, and which can change from moment to moment, dependent on the flying situation. With the aid of the present invention, the pilot is presented directly with the flight path angle and can thus navigate and control the aircraft more readily.
Further in accordance with the present invention, we have developed a simple and reliable method and apparatus for determining the desired fight path and roll angles under arbitrary conditions as long as the flight dynamics are balanced, i.e., the forces required for any centripetal acceleration associated with a maneuver are balanced by the lift and gravitational forces associated with that maneuver, as described more fully below. Specifically, from the data obtained from a measurement system such as a GPS system which provides at least periodic measurements of the position of the aircraft, we obtain velocity and acceleration data from which the desired attitude (flight path angle, roll) information is determined. In accordance with the preferred embodiment, a Kalman filter receives the measurement data and provides the required information from which the desired attitude is determined. In our filter, the aircraft is modeled as a triple integrator system, the filter then providing estimates of the jerk (impulse), velocity, and acceleration of the aircraft during its maneuvers. The latter (velocity and acceleration) estimates provide the requisite information for determination of the desired attitude as described more fully below.
In particular, in accordance with the preferred embodiment of the invention, we determine the flight path angle xcex3 and roll angle xcfx86s of an aircraft and display these as parameters to the pilot as a principal measure of the aircraft attitude at a given moment. The roll angle is advantageously determined about the stability axis of the aircraft, as opposed to the body axis, since determination of the stability axis roll angle requires no knowledge of the angle of attack, a parameter which is frequently not known or readily determinable with accuracy in small aircraft. This is in contrast to conventional aircraft navigation systems which determine and display roll angle and pitch about the aircraft""s body axis. Although the body-axis and stability-axis roll angle are typically nearly equal to each other, this is not the case with pitch angle, which differs from flight path angle by the angle of attack. Thus, when a pilot wishes to maintain the aircraft on a particular flight path angle, he/she can not navigate by the pitch angle alone, but must correct it by the current angle of attack, a value which is frequently known only imprecisely at best in small aircraft, and which can change from moment to moment dependent on the flying situation.
The present invention enables construction of an integrated flight information and control system that presents to the pilot all the information necessary to safely pilot an aircraft but which eliminates costly equipment such as attitude and heading reference systems (AHRS) and other sensors heretofore required to generate the necessary data. In particular, all the essential parameters commonly measured for use in controlling and navigating the aircraft can now be obtained from a single sensor, i.e., a single GPS receiver that provides the requisite outputs from which the position, altitude, vertical speed, ground speed, ground heading, and now, pseudo attitude, of the aircraft can be determined. Thus, an integrated flight information and control system of the type commonly found only in large, commercial aircraft such as a Boeing 747 can now feasibly be provided in small aircraft such as a Piper Arrow. The result is expected to revolutionize aircraft flight information and control systems on such aircraft.
In addition to utilization in small aircraft, the attitude indicator and flight and control system of the present invention is adapted to larger, commercial and military aircraft as well, where it may serve either as the principal navigation and control system or as a backup system or a reference system for integrity checking. Because of the dramatically reduced cost of the system, a significantly greater degree of redundancy may be provided at a reasonable cost, thus further enhancing flight safety overall.